Combined convection/effusion cooled one-piece can combustor

ABSTRACT

An industrial turbine engine comprises a combustion section, an air discharge section downstream of the combustion section, a transition region between the combustion and air discharge section, a combustion transition piece and a sleeve. The transition piece defines an interior space for combusted gas flow. The sleeve surrounds the combustor transition piece so as to form a flow annulus between the sleeve and the transition piece. The sleeve includes a first set of apertures for directing cooling air from compressor discharge air into the flow annulus. The transition piece includes an outer surface bounding the flow annulus and an inner surface bounding the interior surface, and includes a second set of apertures for directing cooling air in the flow annulus to the interior space. Each of the second set of apertures extends from an entry portion on the outer surface to an exit portion on the inner surface.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to means of cooling componentsof a gas turbine, and more particularly, to the cooling of a one-piececan combustor by a combination of convection cooling and effusioncooling.

2. Description of the Related Art

A gas turbine can operate with great efficiency if the turbine inlettemperature can be raised to a maximum. However, the combustion chamber,from which combusted gas originates before entering the turbine inlet,reaches operating temperatures well over 1500° F. and even most advancedalloys cannot withstand such temperatures for extended periods of use.Thus, the performance and longevity of a turbine is highly dependent onthe degree of cooling that can be provided to the turbine componentswhich are exposed to extreme heating conditions.

The general concept of using compressor discharge air to cool turbinecomponents is known in the art. However, developments and variations inturbine designs are not necessarily accompanied by specific structuresthat are implemented with cooling mechanisms for the turbine components.Thus, there is a need to embody cooling mechanisms into newly developedturbine designs.

BRIEF DESCRIPTION OF THE INVENTION

Accordingly, it is an aspect of the present invention to enhanceconventional gas turbines.

To achieve the foregoing and other aspects and in accordance with thepresent invention, an industrial turbine engine is provided thatcomprises a combustion section, an air discharge section downstream ofthe combustion section, a transition region between the combustion andair discharge section, a combustor transition piece defining thecombustion section and transition region, and a sleeve. Said transitionpiece is adapted to carry combusted gas flow to a first stage of theturbine corresponding to the air discharge section. The transition piecedefines an interior space for combusted gas flow. The sleeve surroundsthe combustor transition piece so as to form a flow annulus between thesleeve and the transition piece. Said sleeve includes a first set ofapertures for directing cooling air from compressor discharge air intothe flow annulus. The transition piece includes an outer surfacebounding the flow annulus and an inner surface bounding the interiorsurface. The transition piece includes a second set of apertures fordirecting cooling air in the flow annulus to the interior space. Each ofthe second set of apertures extends from an entry portion on the outersurface to an exit portion on the inner surface.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing and other aspects of the present invention will becomeapparent to those skilled in the art to which the present inventionrelates upon reading the following description with reference to theaccompanying drawings, in which:

FIG. 1 shows an example embodiment of a one-piece can combustor in whichthe present invention can be implemented.

FIG. 2 shows a close-up, perspective view of a sleeve with cooling airentry holes surrounding a transition piece with effusion holes.

FIG. 3 shows a cross-sectional view across the cooling air entry holesof the sleeve and effusion holes of the transition piece.

DETAILED DESCRIPTION OF THE INVENTION

Example embodiments that incorporate one or more aspects of the presentinvention are described and illustrated in the drawings. Theseillustrated examples are not intended to be a limitation on the presentinvention. For example, one or more aspects of the present invention canbe utilized in other embodiments and even other types of devices.

FIG. 1 shows an embodiment of a single piece combustor 10 in which thepresent invention can be implemented. This example embodiment is acan-annular reverse-flow combustor 10 although the invention isapplicable to other types of combustors. The combustor 10 generatesgases needed to drive the rotary motion of a turbine by combusting airand fuel within a confined space and discharging the resultingcombustion gases through a stationary row of vanes. In operation,discharge air from a compressor reverses direction as it passes over theoutside of the combustors 10 and again enters the combustor 10 en routeto the turbine. Compressed air and fuel are burned in the combustionchamber. The combustion gases flow at high velocity into a turbinesection via a transition piece 120. As discharge air flows over theoutside surface of the transition piece 120, it provides convectivecooling to the combustor components.

In FIG. 1, a transition piece 120 transitions directly from a circularcombustor head-end 100 to a turbine annulus sector 102 (corresponding tothe first stage of the turbine indicated at 16) with a single piece. Thesingle piece transition piece 120 may be formed from two halves orseveral components welded or joined together for ease of assembly ormanufacture. A sleeve 129 also transitions directly from the circularcombustor head-end 100 to an aft frame 128 of the transition piece 120with a single piece. The single piece sleeve 129 may be formed from twohalves and welded or joined together for ease of assembly. The jointbetween the sleeve 129 and the aft frame 128 forms a substantiallyclosed end to a cooling annulus 124. It should be noted that “single”also means multiple pieces joined together wherein the joining is by anyappropriate means to join elements, and/or unitary, and/or one-piece,and the like.

In FIG. 1, there is an annular flow of the discharge air that isconvectively processed over the outside surface of the transition piece120. In the example embodiment, the discharge air flows through thesleeve 129 which forms an annular gap so that the flow velocities can besufficiently high to produce high heat transfer coefficients. The sleeve129 surrounds the transition piece 120 forming a flow annulus 124therebetween. As indicated by arrows, cross flow cooling air travelingin the annulus 124 continues to flow upstream in a directionperpendicular to cooling air flowing through holes, slots, openings orother apertures 400 formed about the circumference of the sleeve 129.The sleeve 129 has a series of holes, slots, openings or other apertures400 that allow the discharge air to move into the sleeve 129 atvelocities that properly balance the competing requirements of high heattransfer and low pressure drop. A circled area of the transition piece120 will be discussed in more detail in FIGS. 2-3.

In conventional combustors, a combustor liner and a flow sleeve aregenerally found upstream of the transition piece and the sleeverespectively. However, in the one-piece can combustor of FIG. 1, thecombustor liner and the flow sleeve have been eliminated in order toprovide a combustor of shorter length. The major components in aone-piece can combustor include a circular cap 134, an end cover 136supporting a plurality of fuel nozzles 138, the transition piece 120 andsleeve 129.

FIG. 2 shows a close-up, perspective view of the transition piece 120and the sleeve 129. The sleeve 129 is radially outward with respect tothe transition piece 120 and surrounds the transition piece 120 formingthe flow annulus 124 in between. The sleeve 129 is formed with aplurality of first apertures or holes 400 to allow compressor dischargeair to enter the flow annulus 124 from the exterior space 302. Thesingle-piece transition piece 120 is formed with a plurality of secondapertures or effusion holes 200. It must be noted that FIG. 2 shows oneexample arrangement of the first and second apertures 200, 400 which isnot to be construed as a limitation on the invention. The formation ofthe apertures 200, 400 may be at or extend to other selected areas orover the entire surface of the transition piece 120 and the sleeve 129respectively. The apertures 200, 400 may be formed in acircumferentially dispersed manner or may extend from an upstreamportion to a downstream portion of the transition piece 120 and thesleeve 129 respectively. Moreover, FIG. 2 shows only one of multiplepossible arrangements in which the plurality of apertures 200, 400 canbe patterned. For example, FIG. 2 shows the second apertures 200 inorthogonal arrangement about one another. In another example, eachsecond aperture 200 in a row may be slightly offset relative to secondapertures in an adjacent row. The first apertures 400 are also arrangedin rows and columns but the spacing between the first apertures 400 maydiffer in a row direction relative a column direction. The spacingbetween the first apertures 400 may also differ from the secondapertures 200 as shown in FIG. 3 in part due to the difference in theirsizes. Such variety in arrangement is within the scope of the presentinvention.

FIG. 3 shows a cross-section through the sleeve 129 and the transitionpiece 120. Again, a limited number of apertures 200, 400 are shown onthe transition piece 120 and the sleeve 129 for simplicity ofillustration. In particular, FIG. 3 shows a wall 500 that is part of thesleeve 129 and a wall 300 that is part of the transition piece 120. Thewall 500 separates an exterior space 302 from the flow annulus 124. Thedistance between the wall 300 and the wall 500 may range from 0.5 inchto 3.0 inches.

The first apertures 400 are configured to be normal to the wall 500 suchthat air flow I is adapted to not strike or directly impinge an outersurface 300 a of the transition piece 120 perpendicularly. The firstapertures 400 may be formed directly above the second apertures 200(FIG. 3), may be formed to be offset from the second apertures 200 (FIG.2) so that no second apertures are found below the first apertures 400,or may be formed to be above an area of the wall 300 that in partincludes the second apertures 200 and in part does not include thesecond apertures 200. In a configuration where the first apertures 400are not directly above the second apertures 200, a greater portion ofthe air flow I is allowed to flow over an outer surface 300 a of thetransition piece 120 rather than enter the apertures 200 upon arrival atthe outer surface 300 a.

FIG. 3 also shows an outer surface 300 a and an inner surface 300 b ofthe wall 300. The area above the wall 300 is the flow annulus 124 whilethe area below the wall is the interior space 304 of the transitionpiece 120. A right side of FIG. 3 corresponds to an upstream area withinthe turbine while a left side of FIG. 3 corresponds to a downstream areawithin the turbine. Flow C, made up of compression discharge air whichis cooler than combusted hot gas, originates from the compressor butapproaches the transition piece 120 in the flow annulus 124 from adownstream area of the turbine and moves upstream as is typical in acan-annular, reverse flow combustor. Flow I, also made up of compressordischarge air, moves upstream in the exterior space 302 from adownstream area of the turbine and enters the flow annulus 124 throughthe first apertures 400. Flow H, made up of hot gas, originates from thecombustion chamber and is directed downstream in the interior space 304of the transition piece 120.

As shown in FIG. 3, the second apertures 200 extend from the outersurface 300 a to the inner surface 300 b of the wall 300. The presentinvention encompasses second apertures 200 formed to be normal to thewall 300 and formed at an angle θ to the wall 300. In FIG. 3, theapertures 200 are shown at the angle θ such that exit portions 200 b ofthe apertures 200 are downstream or rearward relative to entry portions200 a of the apertures 200. In one embodiment, the angle θ formed by thelongitudinal axes 200 c of the apertures 200 and a direction 202 that istangential to the wall 300 and is pointed downstream may be acute at 30degrees and may range from 20 to 35 degrees. However, other smaller andlarger angles are also contemplated. In FIG. 3, the downstream tangentpoints to the left. Although the second apertures 200 are substantiallycylindrical, the entry portions 200 a and the exit portions 200 b willhave elliptical shapes if the apertures 200 are not normal to the wall300. However, the apertures 200, 400 may have a cross section that isnot circular and, for example, is polygonal.

Another variation of the apertures 200 is that the angular position ofthe entry portion 200 a may be different from the angular position ofthe exit portion 200 b on the circumference of the transition piece 120.Moreover, the exit portion 200 b of the apertures 200 may be upstream orforward relative to the entry portion 200 a of the apertures 200 therebycreating an obtuse angle between the longitudinal axes of the apertures200 and the direction 202.

In FIG. 3, the second apertures 200 have a substantially cylindricalgeometry with a constant diameter from the entry portion to the exitportion. In one embodiment, the diameter may be 0.03 inch andalternatively may range from 0.02 inch to 0.04 inch. However, otherdimensions for the apertures 200 are also contemplated.

The first apertures 400 also have a substantially cylindrical geometrywith a constant diameter. In one embodiment, the diameter may range from0.1 inch to 1.0 inch. However, other dimensions for the apertures 400are also contemplated.

Also, the apertures 200, 400 may gradually increase or decrease indiameter through the walls 300, 500 respectively.

The second apertures 200 may be formed on the wall 300 of the transitionpiece 120 by laser drilling or other machining methods selected based onfactors such as cost and precision. The larger dimensions of the firstapertures 400 allow for more tolerance and thus similar or morecost-effective machining methods may be used to form the apertures 400.

In FIG. 3, flow I caused by the first apertures or holes 400 cools thetransition piece 120 by forming jets of air that do not strike ordirectly impinge on the outer surface 300 a. Flow C in the flow annulusprovides convective cooling of the transition piece 120 by removing heatwhile traveling along the outer surface 300 a. Flow E created by thesecond apertures or effusion holes 200 provides jets of air at all orselected areas of the transition piece 120 that cool the transitionpiece 120 as the cooling air passes through the apertures 200 contactinginternal surfaces therein. Effusion cooling is a form of transpirationcooling. An aperture that is angled to the wall will have a largerinternal surface area compared to an aperture normal to the wall due toincreased length so that heat transfer is prolonged and greater coolingof the transition piece 120 can be achieved. Moreover, after the coolair exits the exit portion 200 b of the apertures 200, a layer or filmof cooling air is formed adjacent the inner surface 300 b of the wall300 of the transition piece 120. Formation of such a layer of coolingair on the inner surface 300 b further cools the transition piece 120.The formation of such a layer is facilitated by an angled aperturecompared to a normal aperture since the degree of change required indirection by the cool air is reduced. However, the present inventionencompasses the two variations of normal and angled apertures. Coolingby the film formed on the inner surface can improve as the hole sizesand angles are decreased. However, smaller holes are more prone toblockage from impurities. In comparison, larger holes can causeexcessive penetration of the hot gas stream by the cool air jets andreduce the efficiency of the turbine. Therefore, such benefits anddrawbacks must therefore be collectively considered when determining thegeometry of the effusion holes.

The invention has been described with reference to the exampleembodiments described above. Modifications and alterations will occur toothers upon a reading and understanding of this specification. Exampleembodiments incorporating one or more aspects of the invention areintended to include all such modifications and alterations insofar asthey come within the scope of the appended claims.

1. A turbine engine comprising: a combustion section; an air dischargesection downstream of the combustion section; a transition regionbetween the combustion section and air discharge section; a combustortransition piece defining the combustion section and transition region,said transition piece adapted to carry combusted gas flow to a firststage of the turbine engine corresponding to the air discharge section,the transition piece defining an interior space for combusted gas flow;and a sleeve surrounding the combustor transition piece so as to form aflow annulus between the sleeve and the transition piece, said sleeveincluding a first set of apertures for directing cooling air fromcompressor discharge air into the flow annulus, wherein the transitionpiece includes an outer surface bounding the flow annulus and an innersurface bounding the interior space, the transition piece includes asecond set of apertures for directing cooling air in the flow annulus tothe interior space, and each of the second set of apertures extends froman entry portion on the outer surface to an exit portion on the innersurface.
 2. The turbine engine of claim 1, wherein the first set ofapertures are normal to the sleeve.
 3. The turbine engine of claim 1,wherein the first set of apertures has a constant diameter ranging from0.1 inch to 1.0 inch.
 4. The turbine engine of claim 1, wherein one ofthe entry portion and the exit portion is located further downstreamthan the other of the entry portion and the exit portion.
 5. The turbineengine of claim 4, wherein the combustor transition piece is acan-annular, reverse-flow type such that combusted gas flow andcompressor discharge air flow are configured to be in opposingdirections such that longitudinal axes through the second set ofapertures form an acute angle with a direction of combusted gas flow andan obtuse angle with a direction of compressor discharge air flow. 6.The turbine engine of claim 1, wherein longitudinal axes through thesecond set of apertures are oriented to form an acute angle with adownstream tangent to the outer surface.
 7. The turbine engine of claim6, wherein the acute angle ranges from 20° to 35°.
 8. The turbine engineof claim 1, wherein the second set of apertures have a constant diameterfrom the entry portion to the exit portion ranging from 0.02 inch to0.04 inch.
 9. The turbine engine of claim 1, wherein the second set ofapertures are substantially normal to the outer surface.